It is a practice to provide blades or vanes of gas turbines with some form of cooling in order to withstand the high temperatures of the hot gases flowing through such turbines. Typically, cooling ducts are provided within the airfoil of the blades or vanes, which are supplied in operation with pressurised cooling air derived from the compressor part of the gas turbine. Usually, the cooling ducts have the convoluted form of a serpentine, so that there is one flow of cooling fluid or cooling air passing through the airfoil in alternating and opposite directions. However, such a convoluted passageway necessarily requires bends, which give rise to pressure losses without heat transfer. Furthermore, as there is only one flow of cooling fluid, it is difficult to adapt this flow to the various cooling requirements existing at different locations of the airfoil.
To achieve more flexibility in the cooling of the airfoil, it has been described (U.S. Pat. No. 6,874,992) to provide the airfoil with a plurality of cooling passages comprising a plurality of inlet passages along which cooling air flows from the base towards the tip region of the blade and a plurality of return passages along which cooling air flows from the tip towards the base region of the blade, whereby at least some of said inlet and return passages being connected by a common chamber located within the tip region of the blade.
However, as these cooling passages are in fluid communication with each other by means of said common chamber located within the tip region of the blade, it is still difficult to adjust the individual mass flows of cooling fluid flowing through the various cooling passages.
Another problem recognized by the present invention, which is related to the supply of the cooling fluid through the root of the blade or vane, may be explained with reference to FIGS. 1-3:
According to FIG. 1, a blade 10 of a gas turbine comprises an airfoil 14 with a leading edge 17 and a trailing edge 16. The airfoil 14 extends along a longitudinal axis X of said blade between a lower end and a blade tip 15. At the lower end of said airfoil 14, a blade root 12 is provided for being attached to a groove 31 in a rotor 11 of said gas turbine. A hollow blade core 18 is arranged within said airfoil 14 and extends along the longitudinal axis X between said blade root 12 and said blade tip 1. The blade core 18 is provided for the flow of a cooling fluid, which enters said blade core 18 through a blade inlet 20 at said blade root 12 and exits said blade core 18 through at least one dust hole at said blade tip 15. The cooling fluid (cooling air) is supplied by means of a rotor bore 19, which runs through the rotor 11 and is in fluid communication with said blade inlet 20 of said blade 10.
As shown in FIG. 1, the direction of the rotor bore 19 is aligned with the blade orientation, i.e. the longitudinal axis X. A unique passage smoothly distributes the flow all over the cross section of the duct further above the blade inlet 20. However, the area/shape of the rotor bore exit 19, which is cylindrical, and the inlet 20 of the blade, which is race-track shaped, are different, leading to a non-continuous interface (see FIG. 3, the common area is shaded).